(a) Field of Invention
This invention generally relates to a root end portion of a helicopter rotor blade, and more particularly to the root end of a rotor blade fabricated from fiber-reinforced resin materials.
(b) Prior Art
Helicopter rotor blade manufacture has progressed from fabrication from wood spars and ribs and stiffened fabric fairings, to intermediate construction techniques employing bonding discrete, metal spars, ribs and fairing skins into a unitary assembly. More recently, new materials known generally as composites, are being used to form the primary structural components and the secondary non-structural fairings as well. These materials are employed generally in the laminate or roving form and comprise a partially-cured yet flexible resin matrix, usually epoxy, phenolic or similar material, reinforced by continuous thin strands of synthetic fibers. The increasing use of composite materials in the partially cured form, that is later cured to a final shape upon the application of heat and pressure, has permitted the necessary structural reactions to be provided in novel ways that are economic of material and manufacturing labor.
For example, the root end of any rotor blade must be attached to the rotor hub. This attachment must be capable of transferring the dynamic and aerodynamic loads developed within and applied to the rotating blade. Furthermore, the joint must allow easy disassembly of the rotor blade from the rotor hub in a way that requires minimum expertise and standard tools.
Centrifugal force, which is substantially aligned with the spanwise blade pitch axis, and flapwise bending moments are readily introduced to the rotor hub from the rotor blade root end by way of a pinned-joint whose axis is essentially normal both to the blade span and to the rotor mid-plane. However, in-plane bending moments or chord-wise bending is not readily accommodated in a single pin joint whose axis is substantially vertically aligned.
Various arrangements of the structural components in the vicinity of the rotor blade attachment joint have been attempted to produce the requisite strength, stiffness and reliability. In-plane bending is known to produce large and concentrated forces in the rearmost component of the rotor blade known as the trailing edge block. For this reason, struts or rods have been variously attached to the inboard end of the trailing edge at some point that is outboard of the blade-hub attachment joint. The strut, which may incorporate an energy-dissipating damper is attached to the rotor hub at a spanwise position that is inboard of the attachment. In this way, the strut transmits the trailing edge force across the attachment joint and chordwise bending moment continuity is provided. Obviously, this load path requires several individual parts in addition to the principal constituent parts of the rotor blade and involves a structural joint at both ends of the strut. Good engineering practice attempts to limit the number of structural joints, which are well known to present risk of premature failure given the considerable fatigue-intensive nature of the helicopter operating environment.
A second approach taught from preceding developments in the art and particularly directed to root end attachments involving rotor blades formed of composite materials involves use of a machined metal fitting at the attachment joint. Usually this fitting is bonded to the composite materials during or immediately after the heat cycle that cures the composite into its final shape. By way of this bond, the in-plane bending moment is transmitted from the blade root to the fitting for subsequent reaction on the rotor hub. Alternatively, instead of causing the fitting to be bonded to the composite, the metal fitting may be fitted to the outer surface of the blade root in two complementary halves that are clamped shut by bolts. Chordwise moments are transmitted from the blade root to the fitting by bearing contact between the outer root surface and the inner surface of the fitting.
Either use of a machined metal fitting requires a precise matching of the blade root outer contour to the fitting contour. Also, there is inherent disparency in the coefficients of thermal expansion between the metal and the composite blade material that produces residual thermal stress on bonding at elevated temperatures. The additional cost to form the metal part, which usually is of titanium having the associated forming difficulties that are characteristic of this metal, is another disadvantage. Of particular concern where a preformed fitting is incorporated in an assembly having composite material parts is the quality and reliability of the bond that joins the rotor blade structure to the fitting. The mating surfaces must be kept free of contaminants. In addition, the strength of the bond is determined by continuity along its length. Pockets of entrapped air are a common source of interruption in the bond whose presence must be discovered with electronic equipment usually employing sonic detection techniques. Of course, the quality of the bondline is substantiated only after the curing process is complete, but if a substandard joint has been made an elaborate disassembly and rebonding procedure is required.